It is often desirable that an aircraft gas turbine engine include within its compressor, a structure which permits bleeding or diversion of high pressure air from a stage, such as the 5th stage of the compressor to provide high pressure air for cooling purposes and for operation of airframe accessories, engine accessories, or engine or aircraft de-icing systems. In other cases, it is desirable to include a structure which permits the bleeding of even higher pressure air from the discharge of the compressor to provide pressurized air for cooling downstream turbine components. Both interstage bleed and the compressor discharge bleeding are accomplished by flowpath mechanisms which interfere with the normal airflow patterns in the compressor. Further, the casing or bleed structure adds complexity to the assembly of such an engine.
The axial location or stage at which air is bled from the compressor is determined by the pressure required to drive the specific system intended to be serviced by that air. In most instances, it is desirable to achieve the highest possible source pressure to also ensure a high delivery pressure. For this reason, prior systems have extracted air from the latter stages of the compressor and more particularly, engines having these systems have been designed to extract high pressure air from the 5th stage of the compressor for low pressure turbine cooling and turbine thermal clearance control. However, bleeding air from the earliest possible stage of the compressor generally increases compressor efficiency by reducing the amount of work invested in the extracted air. Therefore, it is desirable to achieve the highest possible system supply pressure from the earliest and lowest pressure stage of the compressor. The resulting temperature of the cooling air is also lower and hence more effective.
Known examples of bleed openings or ports can be found in U.S. Pat. No. 4,711,084 to Brockett for an ejector-assisted compressor bleed which discloses a bleed aperture 17 in FIG. 2 having rounded hole edges. U.S. Pat. No. 3,108,767 to Eltis, et al., for a bypass gas turbine engine with an air bleed means in FIG. 3 discloses a duct 19 which is attached to the compressor through a series of chopped holes. U.S. Pat. No. 3,898,799 to Pollert, et al., for a device for bleeding off compressor air in a turbine jet engine, in FIG. 2 discloses a compressor orifice marked with the arrow K. U.S. Pat. No. 3,777,489 to Johnston, et al., discloses a combustor casing having a concentric air bleed structure which includes a series of conical arms 62, 64, and 66 situated in the low velocity area of the diffuser with the bled air structure making a turn of approximately 180.degree.. U.S. Pat. No. 4,344,282 to Anders is directed to a compressor bleed system which includes a locking strap 12 which seals a series of bleed ports 8. U.S. Pat. No. 4,827,713 to Peterson, et al., for a stator valve assembly for rotory machine which includes a passage 30 in the compressor bleed system 28. The structure disclosed in each of these patents significantly reduces the pressure or velocity of the extracted air and thus reduces the energy level of the diffuser air. These documents fail to teach or suggest a pressure efficient diffuser slot which maintains the energy and pressure level of the diffused air to allow the extraction of air from an earlier compressor stage yet having a pressure and energy level equivalent to air previously extracted from a later stage.